1. Field of the Invention
The present invention relates to rocket engines. More particularly, the invention relates to upper stage, restart capable, highly reliable rocket engines.
2. Discussion of the Related Art
Launch vehicles are typically comprised of multiple elements stacked together in stages. There are lower stages and upper stages, with the lower stages being used for lifting the most weight and pushing a vehicle through the thicker layers of lower atmosphere. The upper stages normally fire subsequent to the vehicle traveling very fast in thinner layers of atmosphere and at a very high rate of speed. Accordingly, improvements in upper stage rocket design, even relatively minor improvement can provide a large increase in the performance of a mission.
For example, an improvement in performance for an upper stage can positively affect the amount of payload that a rocket can place into orbit. As a result, some of the most advanced rocket research has focused on upper stages. Although there is no strict definition of an “upper stage,” it usually refers to the second, third, and fourth (if any) stages of a rocket, ignited and fired at high altitude.
Space missions require rocket engines that provide high thrust, high efficiency, and robust durability in order to be operated under the demanding conditions of outer space. For example, there are extreme differences in temperature and pressure both during takeoff and in flight.
There have been many attempts to develop high performance and high reliability rocket engines. One such engine is the standard gas-generator cycle, wherein a small fraction of the overall inlet propellant flow is combusted in a secondary combustion chamber, and the product of this combustion is used to drive the turbomachinery directly. These combustion gases are then discharged. Some functioning or past examples of this engine cycle include the J-2 (Saturn V) and the lower stage engine RS-68 (Delta IV).
A concept closely related but more complex than the gas-generator cycle is called the staged-combustion cycle. In this case rather than discharging the combustion gases from the secondary combustion zone, these gases are introduced into the primary combustion chamber and re-combusted. The result is an engine with extremely high performance but also a high degree of complexity. A functioning example of this type of engine is the main engine of the NASA Space Shuttle.
Yet another typical rocket engine cycle proposed for upper stage application is the tap-off cycle, wherein hot combustion products are tapped off from the primary combustion zone and these gases are used to drive the turbomachinery before being discharged, as in the gas-generator cycle. A functioning example of this type of engine, but one that was never actually realized on a rocket vehicle stage is the J-2S developed by the U.S. during the early 1970's.
Finally, the expander cycle rocket engine, which is well known in the prior art, presents an ideal choice for upper stage use. In general, the fuel is heated before it is combusted, typically with waste heat from the main combustion chamber. The expander cycle is based upon the concept of driving the turbomachinery with gases warmed through regenerative cooling of the thrust chamber assembly, so as to eliminate the need for a secondary combustion zone.
The expander cycle rocket engine has many inherent benefits over other typical cycles such as the standard gas-generator cycle, the tap-off cycle, or the staged combustion cycle. In an expander cycle, the fuel is typically heated before it is combusted, the heat being supplied by waste heat from a main combustion chamber.
FIG. 1 is a simplified illustration of a typical expander engine (closed cycle). The heat from the nozzle 105 and the combustion chamber 101 are used to power a fuel pump 115 and an oxidizer pump 120. As the liquid fuel passes through coolant passages 125 in the walls of the combustion chamber 101, the fuel, typically at supercritical pressures, picks up energy in the form of heat. This energy increase within the coolant passages 125 of the walls of the combustion chamber is sufficient to drive the gas turbine 130. The gas turbine provides the power necessary to drive the fuel pump 115 and the oxidizer pump 120. The turbine discharge along with the pumped oxidizer are then provided to the combustion chamber 101 for combustion.
In a closed-cycle expander engine, the exhaust is sent from the turbine to the combustion chamber, whereas in an open-cycle expander engine only some of the fuel is heated to power the turbines, and then vented, resulting in decreased overall efficiency, though this design does have other potential benefits.
Some disadvantages of all of the aforementioned expander cycle engines includes a limit in the amount of available power to drive the turbomachinery because the driving of the turbomachinery is caused by using the heat extracted through the cooling wall of the primary combustion chamber. Attempting to increase the amount of heat transferred can normally be achieved by using extremely thin walls and/or exotic means of increasing the local wall temperature within the combustion chamber.
However, to design a thinner wall or increase the wall temperature causes a reduction of the structural strength of the combustion chamber wall material, thereby reducing the reliability of the component and the engine. Also, typically such engines are shutdown fuel-rich, meaning that the through cooling passages permit cryogenic hydrogen to flow during the shutdown process. Shutting down the engine fuel-rich is standard practice and is intended to establish a benign environment and to avoid a catastrophic failure. However, when this practice is performed in outer space, the hardware and the passages can stay very cold for a long time. Due to the fact that the expander cycle requires heat to initiate and drive the cycle, restarting the engine in this cold state is difficult and unreliable, as it will take some time before the engine has become warm enough to initiate the start sequence.
FIGS. 2 and 3 show an example of one attempt at improving the expander cycle engine as disclosed in U.S. Pat. No. 6,832,471 to Hewitt. Hewitt discloses that by injecting the oxidizer in two streams, with a smaller stream being injected into the upstream or preburner, and the remainder to the downstream or main combustion section, the use of a cooling element with a high intimate heat exchange construction is permitted to extract a high level of energy from the preburner gas in the form of heat without damaging the cooling element.
More specifically, Hewitt discloses a nozzle 11 of an expander cycle supersonic rocket engine. The drawing shows a combustion chamber 13, a throat 17, a supersonic section or skirt 15. The combustion chamber 13 has an upstream or preburner section 21 and a main downstream section 22. The upstream or preburner section 21 is a secondary combustion zone, which is used instead of the more common gas-generator and a main downstream section. In this configuration, the products from the secondary combustion zone are directly fed directly into the primary combustion zone rather than discharged externally.
Still referring to FIGS. 2 and 3, Hewitt discloses that the first portion of liquid oxygen 23 is fed to an inlet torus 24 that surrounds the upstream portion, wherein the torus directs the liquid oxygen through the chamber wall and into the interior of the chamber. The second portion (remainder) of the liquid oxygen is fed to inlet torus 25 and the heated gaseous hydrogen 27 from the turbopump turbine 41 (FIG. 3) exhaust is fed to the preburner section 21 for combustion. A platelet laminate 31 consisting of a laminate of two stacks of circular disks 33,34, one above the other, separated by a barrier disk 35, is used for heat exchange. The circular disks 33, 34 have central openings 32 and an open space 36 at the stack periphery or by axial channels to form two independent flowpaths that are in heat exchange relationship but not fluid communication. One flowpath is for the combustion gas, and the other is for the uncombusted hydrogen fuel that serves as the coolant.
Still referring to FIGS. 2 and 3, Hewitt discloses the flowpath for combustion gases passes radially outward through the upstream stack 33, then into the annular space 36, then radially inward through the downstream stack 34, then through tubular passages in a distribution manifold 37. The flowpath for the hydrogen fuel acting as a coolant enters the downstream stack 34 upon emerging from the jacket 16, then flows radially inward through the downstream stack 34 (counter-current to the combustion gas) then through axial passages that connect the downstream stack 34 to upstream stack 33, then radially outward through the upstream stack 33 to a space above the upstream stack that leads outwards 38.
The coolant 38 being led outward is now in a gaseous form and directed to the drive turbine 41 of the turbopump (shown in FIG. 3). The heated gas 40 drives turbine 41, which drives two shafts 42, 23 and has separate pumps 44, 45 for liquid hydrogen and fuel. The heated gas 40 pumps fresh coolant to the jacket 16, which flows in the direction indicated by dashed arrows 17,18 while the heated gas itself is expanded and fed 27 to the preburner injector for combustion in the preburner and the main section of the combustion chamber. The partially cooled combustion gas from the preburner 21 is joined by the remainder of the liquid oxygen feed at the downstream face of the injector/manifold 37 to distribute both the fuel-rich preburner gas and freshly supplied oxygen across the width of the combustion chamber. The expanded uncombusted hydrogen 46 that emerges from the drive turbine 41 is then injected into the combustion side of the upstream section 21 of the engine.
Thus, in Hewitt the preburner combustion gas is cooled in a substantially uniform manner to a moderate temperature by cooling the bulk of the gas rather than cooling only the gas in a boundary layer adjacent to the chamber wall.
However Hewitt has disadvantages because the primary and secondary combustion zones are so closely linked. Due to the fact that Hewitt effectively incorporates a staged-combustion element within the expander cycle, this inherently brings into the situation the difficult balancing act of the interplay between the two combustion zones. Hewitt also suffers from some of the disadvantages of the other engines as well.
Other drawbacks of any of the above-mentioned rocket engines, including Hewitt, include the problem that after the engine is shut down, the residual combustion products, specifically steam, can freeze in the injectors and cause damage. Moreover, for those engine cycles using combustion products as the turbine drive gas, the steam can also freeze in the turbomachinery, which has the potential to be even more disastrous upon attempted engine restart than the potential damage from steam freezing in the injectors. Thus, there is a need in the art for an improved expander cycle rocket engine.